Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine vane and blade assemblies to these high temperatures. As a result, turbine vanes and blades must be made of materials capable of withstanding such high temperatures. In addition, turbine vanes and blades often contain cooling systems for prolonging the life of the vanes and blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine airfoils are formed from an elongated portion having a tip at one end and a root coupled to a platform at an opposite end of the airfoil. The root is configured to be coupled to a disc. The airfoil is ordinarily composed of a leading edge, a trailing edge, a suction side, and a pressure side. The inner aspects of most turbine airfoils typically contain an intricate maze of cooling circuits forming a cooling system. The cooling circuits in the airfoils receive air from the compressor of the turbine engine and pass the air through film cooling channels throughout the airfoil. The cooling circuits often include multiple flow paths that are designed to maintain all aspects of the turbine airfoil at a relatively uniform temperature. At least some of the air passing through these cooling circuits is exhausted through orifices in the leading edge, trailing edge, suction side, and pressure side of the airfoil.
Many conventional turbine airfoils have cooling channels positioned at the leading and trailing edges and the outer walls. The airfoils often have a mid-chord cooling channel that may have a serpentine configuration or other design. Often times, the cooling channel is pressurized with cooling fluids to provide adequate cooling fluids to all portions of the cooling channels forming the cooling system in the airfoil. The walls forming the pressurized mid-chord cooling channel often remain at temperatures much lower than other portions of the airfoil in contact with hot combustion gases, thereby resulting in a large thermal gradient between these regions. The large thermal gradient often results in a reduced mechanical life cycle of airfoil components and poor thermal mechanical fatigue (TMF). Therefore, the inner cooling channel often negatively affects the life cycle of the airfoil. Thus, a need exists for a turbine airfoil having increased cooling efficiency for dissipating heat while reducing the thermal gradient between the cooling channels and the hot combustion gases.
FIG. 1 shows an external pressure profile for an airfoil. For a conventional five pass serpentine mid-chord cooling channel, cooling fluid is discharged on the pressure and suction sides. The pressure side of the airfoil has a higher external pressure than the suction side, and thereby is used to determine the pressure of the cooling fluid within the cooling system of the airfoil. In order to meet back flow margin criteria, a high cooling supply pressure is needed for this particular design, which results in a large leakage flow of cooling fluids. The second, third, and fourth passes of the serpentine cooling channel typically include film cooling holes for both the pressure and suction sides. In order to meet the back flow margin criteria for the pressure side film cooling holes, the pressure of the cooling fluid within the serpentine cooling channel is approximately ten percent higher than the pressure side external hot gas pressure. This results in over-pressurizing the suction side film cooling holes, which results in tremendous cooling system inefficiencies.